This invention relates to combustors used in gas turbine engines, and specifically to combustors having ceramic matrix combustor liners that can interface with engine components made from different materials having dissimilar thermal responses.
Improvements in manufacturing technology and materials are the keys to increased performance and reduced costs for many articles. As an example, continuing and often interrelated improvements in processes and materials have resulted in major increases in the performance of aircraft gas turbine engines. One of the most demanding applications for materials can be found in the components used in aircraft jet engines. The engine can be made more efficient resulting in lower specific fuel consumption while emitting lower emissions by operating at higher temperatures. Among the current critical limitations on the achievable operating temperatures of the engine are the materials used in the hottest regions of the engine, which include the combustor portion of the engine and the portions of the engine aft of the combustor portion including the turbine portion of the engine. Temperatures in the combustor portion of the engine can approach 3500xc2x0 F., while materials used for combustor components can withstand temperatures in the range of 2200-2300xc2x0 F. Thus, improvements in the high temperature capabilities of materials designed for use in aircraft engines can result in improvements in the operational capabilities of the engine.
One of the portions of the engine in which a higher operating temperature is desired so that overall operating temperature of the engine can be achieved is the combustor chamber. Here, fuel is mixed with air and ignited, and the products of combustion are utilized to power the engine. The combustor chambers include a number of critical components, including but not limited to the swirler/dome assembly, seals and liners. In the past, these components have been made of metals having similar thermal expansion behavior, and temperature improvements have been accomplished by utilization of coatings, cooling techniques and combinations thereof. However, as the operating temperatures have continued to increase, it has been desirable to substitute materials with higher temperature capabilities for the metals. However, such substitutions, even though desirable, have not always been feasible. For example, as noted previously, the combustors operate at different temperatures throughout the operating envelope of the engine. Thus, when differing materials are used in adjacent components of the combustor, or even in components adjacent to the combustor, widely disparate coefficients of thermal expansion in these components can result in a shortening of the life cycle of the components as a result of thermally induced stresses, particularly when there are rapid temperature fluctuations which can also result in thermal shock.
The concept of using non-traditional high temperature materials such as ceramic matrix composites as structural components in gas turbine engines is not novel. U.S. Pat. Nos. 5,488,017 issued Jan. 30, 1996 and U.S. Pat. No. 5,601,674 issued Feb. 11, 1997, assigned to the assignee of the present application, sets forth a method for making engine components, of ceramic matrix components. However, the disclosure fails to address problems that can be associated with mating parts having differing thermal expansion properties.
U.S. Pat. Nos. 5,291,732 issued Mar. 8, 1994, U.S. Pat. No. 5,291,733 issued Mar. 8, 1994 and U.S. Pat. No. 5,285,632 issued Feb. 15, 1994, assigned to the assignee of the present invention, address the problem of differential thermal expansion between ceramic matrix composite combustor liners and mating components. This arrangement utilizes a mounting assembly having a supporting flange with a plurality of circumferentially spaced supporting holes. An annular liner also having a plurality of circumferentially spaced mounting holes is disposed coaxially with the flange. The liner is attached to the flange by pins that are aligned through the supporting holes on the flange and through the mounting holes on the liner. The arrangement of the pins in the mounting holes permits unrestrained differential thermal movement of the liner relative to the flange.
The present invention provides an alternate arrangement for reducing or eliminating thermally induced stresses in combustion liners and mating parts while permitting unrestrained thermal expansion and contraction of combustor liners.
The present invention provides for a combustor having liners made from ceramic matrix composite materials (CMC""s) that are capable of withstanding higher temperatures than metallic liners. The ceramic matrix composite liners are used in conjunction with mating components that are manufactured from metallic materials. To permit the use of a combustor having liners made from CMC materials in conjunction with metallic materials used for the mating forward cowls and aft seals with attached seal retainer over the broad range of temperatures of a combustor, the combustor is manufactured in a manner to allow for the differential thermal expansion of the differing materials at their interfaces in a manner that does not introduce stresses into the liner as a result of thermal expansion.
A significant advantage of the present invention is that the interface design that permits the differential thermal expansion of the various materials of the components permits the use of ceramic matrix composites for combustor liners by eliminating the thermal stresses that typically shorten the life of the combustors as a result of differential thermal expansion of the parts. The use of the CMC liners allows the combustors to operate at higher temperatures with less cooling air than is required for conventional metallic liners. The higher temperature of operation results in a reduction of NOX emissions by reducing the amount of unburned air from the combustor.
A second advantage of the combustor of the present invention is that is addresses the problems associated with differential thermal growth of interfacing parts of different materials.
Yet another advantage of the present invention is that the interface connections between the CMC liners and the liner dome supports regulates part of the cooling air flow through the interface joint to initiate liner film cooling. Thus, cooling air flow across the combustor liner is not solely dependent on cooling holes as in prior art combustors and state-of-the-art CMC manufacturing technology can be used to manufacture the liners.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.